Method of Manufacturing Aircraft Engine Parts Utilizing Reusable And Reconfigurable Smart Memory Polymer Mandrel

ABSTRACT

A method for fabricating aircraft engine external target parts including complex geometries utilizes reusable reconfigurable shape memory polymer and conformable woven braided carbon fiber sleeves. The method includes providing a tubular three-dimensional reusable shape memory polymer mandrel assembly designed for a target part, and heating the shape memory polymer mandrel.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a divisional of U.S. patent application Ser. No.15/337,931, filed Oct. 28, 2016, which is incorporated by referenceherein in its entirety.

FIELD OF THE INVENTION

The present invention relates generally to duct manufacturing processes.More particularly, the present disclosure relates to fabrication ofmetal and non-metal parts for aircraft engines.

BACKGROUND OF THE INVENTION

In modern aircraft engines, various tubes and ducts which are used todeliver a variety of fluids (e.g., air, oils, fuels, etc.) are generallyknown as fluid delivery systems. These tubes and ducts can have complexgeometry in three-dimensional (3-D) space. For example, the ducts canhave: multiple bends, cylindrical and non-cylindrical geometries,non-uniform cross-sectional size (tapering) and other variable crosssections along an axial direction, and complex transitions fromcylindrical to rectangular cross-sectional shapes. Such designs bringmanufacturing challenges and problems.

Solutions for fabricating metal parts, for example, often involve aseries of operations of bending, stamping, forming, welding, and/orbrazing. In addition, the sub-components may have to be made thickerthan required to compensate for thermal deformation introduced by thejoining process. Moreover, if a design change is made, a new set oftooling is needed. For a final tube or duct assembly, a number ofsub-components may be joined together by fusion welding or brazing.

Solutions for fabricating non-metal parts, for example, often use arigid cure tool/mandrel. Removal of such a mandrel from the duct orcured part, however, can be difficult, costly, and time consuming. Thisis especially so if the duct, or formed part, has a complex geometry(such as those often found in aircraft engine parts) which precludeseasy removal of the part from the rigid mandrel.

Other known manufacturing processes use a segmented mandrel which can beremoved and disassembled in sections, after the formed part has beencured into a rigid state. These types of mandrels, however, can beexpensive and time consuming to install and remove. Moreover, suchsegmented mandrels are often designed to make a single specific part,and are not easily reconfigured to form other different types of parts.

Still other known methods of removing the mandrel involve sacrificingthe mandrel by cutting, dissolving or otherwise breaking the mandreldown into more easily removable pieces. For example, Hammer et al,published Patent Application No. US2014/0023812A1, teaches removing amandrel, used for making aircraft fluid delivery pipes, by dissolvingthe mandrel in acid. Such destructive methods typically not only preventreuse of the mandrel, but may also result in damage to the duct orformed part itself. Moreover, use of expendable tools can involve notonly recurring tooling costs but additional cost for piece work partprocessing as well.

Still other known production processes may use less destructiveinflatable/deflatable balloon-like mandrels. These bladder-type tools,however, often suffer from a lack of strength and rigidity, and mayrequire supporting struts or other load support structures, during thepart fabrication process. Such collapsible tooling, however, is alsocomplex and cost prohibitive for most designs.

Other conventional techniques include making aircraft engine parts usinga substrate, model, mold or mandrel in conjunction with a toolingassembly. A mold, set of molds, or set of pieces to make a mold issometimes referred to simply as a tool or tooling.

For example, making composite parts using shape memory polymer (SMP)mandrels is a well-known process. See U.S. Pat. No. 7,422,714B1, for a“Method of Using Shape Memory Material For Composite PartManufacturing”, issued to Hood et al, on Sep. 9, 2008; and see Methodsand systems for co-bonding or co-curing composite parts using arigid/malleable SMP apparatus, issued to Havens et al, on Feb. 10, 2015.

SMPs are polymeric materials whose qualities are altered to enabledynamic shape “memory” properties. SMPs derive their name from theirinherent ability to return to their original “memorized” shape afterundergoing a shape deformation. SMPs (that have been pre-formed) can bedeformed to any desired target shape. After deformation, in order tomaintain the thermoformed shape (i.e., to “lock” in the deformation),the SMP must remain below, or be quenched to below “Tg”. (Tg is theinherent glass transition temperature of the particular type of SMPselected to be utilized for deformation.) The SMP will hold its deformedshape indefinitely until heated above its Tg again; whereat the SMPreturns to its pre-formed (“remembered”) state. Such SMP properties areknown in the art; (see, for example, Hood′ 714 at column 3).

Although the Hood′714 method employs reusable SMP mandrel, itcontemplates that the SMP is in the form of particle, foam or gel,and/or that the SMP is reinforced with fibrous material (such as textilefabric

SUMMARY OF EMBODIMENTS OF THE INVENTION

Given the aforementioned deficiencies, a need exists for an improvedmethod for forming engine parts, that does not suffer from the abovelimitations, solves some of the concomitant problems indicated above,and employs tooling with mold/mandrel materials that enable both reuseand reconfiguration. A need also exists for an alternate method tofabricate tube and duct assemblies for fluid delivery in aircraftengines.

An embodiment of the present invention includes a method using lowtemperature deposition of metal onto reusable reconfigurable SMPtooling. An embodiment of the present manufacturing method conceptuallyincludes three stages, as discussed in further detail below.

Under certain circumstances, an embodiment of the present inventionincludes a method for fabricating a target part for fluid delivery usingreusable reconfigurable SMP and low temperature metal deposition. Themethod includes providing a mandrel using SMP having a pre-form shapeand a glass transition temperature (Tg), providing a tooling assemblywith a shape mold designed in accordance with the target part andplacing the pre-form SMP in contact with the shape mold and stimulatingthe SMP until the SMP deforms to replicate a geometry of the shape mold.The method also includes coating an outer surface of the deformed SMPwith a conductive metal and placing a whole tooling into a lowtemperature electroplating tank, applying voltage to electrochemicallydeposit metal onto a surface of the tooling, removing the whole toolingfrom the bath and applying heat thereto, the SMP becoming malleable inresponse to the heat.

In addition to use in making metal internal aircraft engine parts,another embodiment of the present invention can be used in fabricatingnon-metal external aircraft engine parts. These parts may be found, forexample, on the exterior of turbine engines. Another exemplaryembodiment uses SMP mandrels, in conjunction with resin coated wovenconformable carbon fiber sleeves, to fabricate the non-metal externalparts.

Still other embodiments could be used, for example, not only for lowtemperature tubes and ducts for some fluid delivery systems, but also tomake composite laminate parts for mounting brackets, fan case housings,supports, and other aerospace components.

Many of the shape shifting properties of SMP materials have been knownin the art for decades, (e.g., see the above discussed Hood'714 patent,and the other SMP related references cited in those patents).Nevertheless, known processing techniques or methods do not teach theconcepts described herein. Known processes certainly do not teachtechniques or methods for fabricating aircraft engine parts havingcomplex 3-D tube and duct geometries with transitioning cross-sections,using smart polymer tooling in the novel fashion disclosed and claimedin the present invention.

Further features and advantages of the invention, as well as thestructure and operation of various embodiments of the invention, aredescribed in detail below with reference to the accompanying drawings.It is noted that the invention is not limited to the specificembodiments described herein. Such embodiments are presented herein forillustrative purposes only. Additional embodiments will be apparent topersons skilled in the relevant art(s) based on the teachings containedherein.

BRIEF DESCRIPTION OF THE DRAWINGS

The accompanying drawings, which are incorporated herein and form partof the specification, illustrate the present invention and, togetherwith the description, further serve to explain the principles of theinvention and to enable a person skilled in the relevant art(s) to makeand use the invention.

FIG. 1. illustrates an exemplary aircraft engine duct in accordance withvarious aspects described herein.

FIG. 2. is a conceptual overview of a manufacturing process, for ductsand tubes, in accordance with an embodiment of the present invention.

FIG. 3 is a flowchart illustrating steps of an exemplary first stage(i.e., making the mandrel) of the process of FIG. 2, in accordance withthe embodiments.

FIG. 4 is a flowchart illustrating steps of an exemplary second stage(i.e., metal deposition) of the process of FIG. 1.

FIG. 5 is a flowchart illustrating steps of an exemplary third stage(i.e., removal and reuse) of the process of FIG. 1.

FIG. 6 illustrates steps of an exemplary high-level method of practicingan embodiment of the present invention.

FIG. 7 depicts a section of an aircraft turbine engine with externalparts to which the embodiments can apply

FIG. 8 is a flowchart of an exemplary method of an embodiment of thepresent invention for fabricating carbon composite external engineparts.

DETAILED DESCRIPTION OF EMBODIMENTS OF THE INVENTION

The following detailed description of the invention references theaccompanying drawings that illustrate specific embodiments in which theinvention can be practiced. The embodiments are intended to describeaspects of the invention in sufficient detail to enable those skilled inthe art to practice the invention. Other embodiments can be utilized andchanges can be made without departing from the scope of the presentinvention. The following detailed description is, therefore, not to betaken in a limiting sense. The scope of the present invention is definedonly by the appended claims, along with the full scope of equivalents towhich such claims are entitled.

In this description, references to “one embodiment”, “an embodiment”, or“embodiments” mean that the feature or features being referred to areincluded in at least one embodiment of the technology. Separatereferences to “one embodiment”, “an embodiment”, or “embodiments” inthis description do not necessarily refer to the same embodiment and arealso not mutually exclusive unless so stated and/or except as will bereadily apparent to those skilled in the art from the description. Forexample, a feature, structure, act, etc. described in one embodiment mayalso be included in other embodiments, but is not necessarily included.Thus, the present technology can include a variety of combinationsand/or integrations of the embodiments described herein.

FIG. 1 is an illustration of an exemplary aircraft engine duct assembly100, having various subcomponents, to which embodiments of the presentinvention may apply. Moreover, as shown in FIG. 1, the duct assembly 100has complex geometries in 3-D space and includes: multiple bends 102,106; a flex joint 104; and variable cross sections 108, 110, 112, and114 along an axial direction. Although the engine part 100 is a duct,the embodiments can be applied to a variety of tubes, assemblies, andaircraft fluid delivery channels, as understood by those of skill in theart.

FIG. 2 is a high-level overview of a multi-stage process 200 of anexemplary embodiment. Each of the multiple stages of the process 200 arediscussed in greater detail below. At a high-level, the process 200 useslow temperature metal depositions, onto an SMP tooling mandrel, tofabricate aircraft engine parts with complex geometries. These engineparts may have complex geometries, such as those discussed above withreference to FIG. 1.

In the multi-stage process 200, before metallization a tooling mandrel,which is part of a tooling assembly, is made of SMP at Stage 1. Understimuli, the SMP can exhibit a change from a rigid state to a verypliable elastic state. While pliable, the SMP can be triggered intomimicking or conforming to other shapes, without degradation.

At Stage 2, and after the mandrel has replicated or conformed to thedesired shape of a target part (hereinafter Part-A), low temperaturemetal deposition can be used to affix a metal layer to an outer surfaceof the SMP tooling. The surface of the tooling is then made electricallyconductive and the entire tooling assembly is put into an electroplatingtank or bath. The metal is electrochemically deposited onto the surfaceof the tooling with application of appropriate voltage. At Stage 2B,additional metal layers may be deposited, as desired, and/or whennecessary.

Once a desired metal deposition is completed, the assembly is heated toa temperature above the glass transition temperature (Tg) of theselected SMP. As a result, at Stage 3, the SMP tooling becomes soft andcan be removed and separated from the newly formed Part-A Removal canoccur without damage to the tooling or to Part-A. If the stimuli arethen removed, the SMP returns to a rigid memory state, and can be usedfor another part.

In the embodiments, pre-machined flanges can be attached to the mandrelbefore electroplating. Metal can then be electroformed over a base ofthe flange and the surface of the mandrel. When the mandrel is removed,Part-A is formed with the flange attached as a subcomponent.

In the description of FIGS. 3, 4, and 5; Stages 1, 2, and 3 will bediscussed in greater detail. In particular, FIG. 3 corresponds to Stage1, FIG. 4 corresponds to Stage 2, and FIG. 5 corresponds to Stage 3.

Stage 1: Making SMP Mandrel

FIG. 3 is a flowchart illustrating more detailed steps of the Stage 1(i.e., making the mandrel) of the process of FIG. 2.

In this exemplary method, starting at step 310, a specific type pf SMPis selected or prepared based on chemical, electrical, thermal and othercharacteristics of particular polymers and the target part in step 315.Numerous types of polymers and alloys, which exhibit shape memoryproperties, are known by those skilled in the art. For example, see theabove cited Hood'714 patent, at Columns 3 and 5; or see Havens′375, atColumns 5 and 6.

The SMP utilized may also be selected or tailored to meet the specifictolerances (e.g., for Part-A) and temperature requirements needed (e.g.,for the selected type of electroforming). One requirement, for examplefor metal parts, is that the SMP utilized, to later form the mandrel,should be selected such that it has a predetermined Tg higher than thetemperature of the deposition bath (which is to be used later in themetal embodiments). For both metal and non-metal embodiments, Tg is amajor consideration. Cost is another consideration in selecting aparticular type of SMP material.

At step 320, a tooling assembly, or tooling, is designed having thedesired shape and geometries of a particular aircraft engine fluiddelivery channel, duct or tube. The tooling is made for the complexgeometry of a target part, such as Part-A, having plural subcomponents.

Prior to contact with the tooling 320, the selected SMP material 315 istypically in a rigid pre-form state. While the embodiment of FIG. 2, forillustrative purposes, indicates an SMP tooling having an annular orcylindrical pre-form geometry, other pre-form memory shapes (such as:spheres and ovals, triangles and pyramids, rectangles and boxes ofvarious shapes and sizes, and other geometric patterns) are within thescope of the present invention.

At step 330, predetermined stimuli, which can later trigger deformationof the selected SMP, are selected and/or calculated. By way of example,depending upon the SMP material selected, the stimuli can be thermal orother stimuli.

At step 325, depending upon the shape of the target part, selected SMPmaterial can be placed onto or inside the mold or tooling assembly. Forexample, in some embodiments of the invention, a substantially tubularor cylindrical pre-form SMP mandrel is placed inside a clamshell mold ormold cavity of a tooling of the desired shape.

Selected stimuli are applied to the tooling to trigger a state change inthe SMP, as indicated in step 335. In some embodiments, the stimulationstep could include use of pressurization, vacuum, and/or various heatedmaterials, including air, gas or other fluids, to facilitate raising thetemperature of the utilized SMP above the material's Tg. In oneembodiment, the mold is put in an oven, which is heated above Tg. Whilein the oven, the SMP is pressurized by air or another gas.

Stimulation of the SMP tooling continues (steps 340, 350, 355, 360)until the SMP temperature exceeds Tg. In steps 350 and 365, the SMPtransforms from a rigid substance into an elastic, flexible, and softsubstance. In this soft malleable elastic state, the SMP mimics orreplicates the tooling mold shape, and thus defines the desired targetgeometry of Part-A.

At step 370, the SMP mandrel is kept within the shape mold and cooled toa temperature below its Tg. It is thereafter removed from the mold. TheSMP is desirably maintained at low temperature to prevent softening andloss of its geometry accuracy. At step 390, of the Stage 1, the process200 transitions to Stage 2 for metal deposition.

Stage 2: Depositing Metal

During Stage 2, metals are deposited onto the outer surface of the SMPtooling. Those skilled in the art would recognize that metal depositionon plastic has been demonstrated previously. Nevertheless, this metaldeposition desirably occurs in a low temperature process to prevent thetooling will become soft and lose its geometric accuracy. Those skilledin the art would know that the Tg of SMP is typically much higher thanmost electroplating processes. Therefore, during the deposition steps,the temperature of the tooling (Tt) is kept well below the melting point(Tm) of the metal to be deposited on the mandrel. Preferably, Tt is kept(<0.5Tm) of the utilized metal. In other words, [Tt<0.5Tm]. As a result,thermal deformation to the assembly is minimized, and thinner andlighter assembly is possible.

By way of example, and not limitation, various deposition methods usedin different embodiments of the present invitation can includeelectroforming, such as by electroplating. Other deposition methodsinclude kinetic metal deposition (cold spray) or ultrasonicconsolidation. In a preferred embodiment, deposition of metal ontoplastic-type materials is achieved, at low temperature, usingelectroplating. In part, this is because the Tg of SMP is typically muchhigher than the temperature of most electroplating processes.

Common materials for electroforming and electroplating are limited to afew, including copper alloy, nickel alloy, aluminum alloy, titaniumalloy and gold. Other common engineering materials (such as stainlesssteel and titanium) cannot always be electrochemically deposited in thedesired fashion. In such cases, a secondary low-temperature metaldeposition technique is required, such as kinetic metal deposition(e.g., cold spray) or ultrasonic consolidation. Various metal stocks canbe used for multi-material parts.

For example, we can achieve layered copper-nickel material, byelectroplating copper to a certain thickness and then electroplatingnickel on top of the copper. The benefit of such material is toleverage, for example, the good thermal conductivity of copper, and usethe outer layer of nickel as a protective layer.

FIG. 4 illustrates the steps of Stage 2 of the present invention,beginning at method step 410. At step 415, an electrically conductivemetallic material is selected and/or prepared. At step 420, the surfaceof the tooling is made electrically conductive by coating it with theconductive material. This can be achieved by painting, spraying, dipcoating, sputtering, electroplating, or other methods. In step 430, thetooling is put into an electroplating bath or tank, and the lowtemperature electroplating process is started at step 435. As voltage isapplied, metal is electrochemically deposited on to the surface of thetooling, at step 440.

Thickness, chemical composition, and/or metallurgical microstructure ofthe metal can be controlled (step 445) by changing voltagecharacteristics, chemical composition, or deposition time. Throughoutstage 220, [Tt<0.5Tm] is maintained (step 460), as one or more layers ofmetal are deposited or applied to the tooling.

As a precursor, the desired number of layers (N) is calculated (425) orset. For example, N may be based on the desired strength or thickness ofPart-A, on mechanical or chemical properties of the selected metal, oron other factors. The first layer (N=1) deposited, on the tooling outersurface, has to be electroplated.

On the other hand, after the first layer is electroplated, furtherlayers (N=>2), could be electroplated, cold sprayed, or ultrasonicallyconsolidated. We could, for example, electroplate layer 1 as material A,electroplate layer 2 as material B, electroplate layer 3 as material C,so on and so forth. Alternatively, we could electroplate layer 1 asmaterial A; and then materials B, C, etc. could be layered on by coldspraying or ultrasonic consolidation.

Electroplating continues (steps 470 and 475) until N layers of metalhave been deposited, at step 480. Once the desired layers for Part-Ahave been applied to the tooling, electroplating is discontinued anddeposition stops at step 485. The tooling assembly, including the metalcoated SMP mandrel, is then removed from the bath at step 490, forseparation in Stage 3 of the instant fabrication method.

Stage 3: Separation & Reuse

FIG. 5 illustrates Stage 3 of the present invention, starting at step510. At step 515, the entire tooling assembly is heated until thetemperature of the tooling exceeds the SMP transition temperature[Tt>Tg]. Heating continues (steps 520 and 525) until the SMP toolingbecomes soft, at step 530.

Thus, at step 530, the tooling mandrel can be easily separated orremoved, from the newly formed Part-A, without damage to either themandrel or to Part-A. Thereafter, in some cases, a post-processing step540 of the new duct may be required for surface conditioning ortolerance of Part-A.

At step 540, the fabricators have the option, if continuing a productionrun of Parts A is desired, of returning to Stage 1 (e.g., at step 320)and fabricating another identical Part-A. If continuing the productionone is not desired at step 555, the newly formed Part-A can be stored orused, at step 560, in an engine fabrication or repair process.

In addition, fabricators have the option to reuse the tooling assemblyfor a different target part (steps 570, 575). If a need exists to make adifferent part (e.g., Part-B), fabricators can go to Stage 1 andreconfigure the tooling assembly to make a Part-B (e.g., by makingappropriate adjustments at steps 315, 325 and/or 330).

Alternatively, at step 580, the tooling assembly can be cleaned andstored (step 585) for later reconfiguration.

FIG. 6 illustrates formation of flanges for attachment to the duct(Part-A) as part of the electroforming process, in accordance with theembodiments. Pre-machined flanges 610 can be attached the SMP mandrel620 during the Stage 1, tooling design step 320. Flanges may also beattached to the SMP mandrel at the start of Stage 2 (see step 410).

The entire tooling, including the attached flange, can be put into theelectroplating tank (e.g., step 420). Metal formation then proceeds byelectroforming 630 over the external surface of the mandrel 620, asillustrated at step 460.

Thereafter in Stage 3 (step 530), the newly formed Part-A is separatedfrom the tooling and the mandrel is removed. The metal coated flange,however, remains attached to Part-A. Thus, the flange can thereby becomea subcomponent of the newly formed duct (Part-A) 640.

Non-Metal Aspects

With the continued development of high temperature resins for compositeaerospace components, there has been an increased use of composites inturbine engine externals. These include carbon composite laminatemounting brackets, fan case housings, and supports. For example, seeFIG. 7, depicting an aircraft engine turbine 700, with an external part702.

Low temperature tubes and ducts are potentially another materialreplacement opportunity for components traditionally made from straightcylindrical stainless steel tubes. These tube typically are bent andformed with hydraulic mandrel tools and checked in gauge fixtures. Thefinal formed tube is often a complex 3-D multiple-bend component. Toproduce a similar lightweight component out of carbon composites canrequire complex multiple degree of freedom robotic tooling for thecombined mandrel positioning and material application. This willsignificantly increase both the initial manufacturing investment andprocessing time.

Complex tooling is often a key to production of both internal andexternal aircraft engine parts; but such increased complexity mayrequire long lead times and is often expensive. For example, molds for acomponent part with only one 60 degree angle can have dramaticallyhigher cost than that of a simple pipe or tube. Thus, workers in theaircraft liquid delivery tube art have recognized the need for othersolutions for manufacturing complex geometry multi-bend multi-radiiducts.

For example, see Hammer′812, discussed above. regarding use of alternateduct fabrication materials. Hammer, however, contemplates use of 3-Dprinting technology with dissolvable mandrels. Also, Hammer expresslyteaches away from use of smart memory materials (see Hammer′ 812 at page2., Column 2, paragraphs 26, 28) as used in the embodiments.

Attempts to post bend straight woven carbon composite tubes have alsobeen made. These have included local inductively heated and resistanceheated die tests, which were not successful, and showed little potentialfor future success. On the other hand, an embodiment of the presentinvention uniquely solves this problem by forming the complex geometrytubes before a thermoset resin is applied to braided carbon fibers andcured. The instant unique conformable woven composite sleeve approach,used with smart tooling, is a simple solution to a challengingmanufacturing process (that usually requires complex multiple degree offreedom robotic equipment and tooling).

An adaption of the present invention, for composite ducts, utilizesconformable woven braided carbon fiber sleeves, along with hightemperature thermoset resins (BMI, benzoxazine, phthalonitrile, orother). This adaptation also utilizes reusable SMP mold/tool toeffectuate a simplified fabrication design solution for non-metalexternal parts of aircraft turbines, other turbines or engines, andother aerospace components.

FIG. 8 is a flowchart 800 of an exemplary method of an embodiment of thepresent invention for fabricating carbon composite external engineparts.

Method step 810 shows the starting point for an embodiment of thepresent invention as used to fabricate non-metal parts for aircraftturbine engines. In ensuing steps, starting materials, used in theinventive process, are preliminarily selected or prepared usingtechniques known to those skilled in the art.

At step 820, a tubular 3-D reusable smart-polymer tooling mandrel isprovided. Also at this stage, a tooling shape mold is provided, in amanner discussed earlier. At step 830, properly sized conformable wovenbraided carbon fiber sleeves are provided. At step 840, a hightemperature resin system, formulated for carbon fibers, is provided.

At step 850, a first layer of properly sized braided carbon fibersleeve, is applied to the SMP mandrel, using a high temperature resinsystem for wetting and adhesion. The number of subsequent layers (N)required may be determined, for example, based on the loads and boundaryconditions for the target engine part (Part-C). Also, multiple localproperly orientated braided sleeves may need to be applied to increasedirection strength at high stress locations. Once it is determined thatsufficient layers have been applied (steps 860 and 865) to the SMPmandrel, overlaying ceases at step 880.

The now carbon fiber coated SMP mandrel, which has been covered with arelease agent (step 825), is heated at step 875 until the Tt exceeds theSMP Tg. Accordingly, the mandrel becomes compliant and pliable. At step880, the elastic SMP can be easily removed and separated from the newlyformed carbon composite Part-C.

Optionally, after removal of the SMP, if needed (step 885) for aparticular type of target part, a subsequent higher temperature resincure may be performed on Part-C, at step 890.

At step 895, the SMP plastic is cleaned, heated, and optionally placedin the tooling shape mold again, for re-forming and reuse. Thus ends thecycle for one embodiment of the inventive fabrication method.

CONCLUSION

While the present invention is described herein with illustrativeembodiments for particular applications, it should be understood thatthe invention is not limited thereto. Those skilled in the art withaccess to the teachings provided herein will recognize additionalmodifications, applications, and embodiments within the scope thereofand additional fields in which the invention would be of significantutility.

The present invention has been described above with the aid offunctional building blocks illustrating the implementation of specifiedfunctions and relationships thereof. The boundaries of these functionalbuilding blocks have been arbitrarily defined herein for the convenienceof the description. Alternate boundaries can be defined so long as thespecified functions and relationships thereof are appropriatelyperformed.

The inventive method utilizes a reusable and reconfigurable tooling,which will be cheaper, more versatile and faster to change, than toolingused in relevant prior art metal part fabrication processes. Multiplesub-components can be integrated into a single one without increasingmanufacturing complexity.

The preferred metal deposition embodiment is conducted at lowtemperature (<0.5Tm of the metal). As a result, thermal deformation tothe assembly is minimized, and thinner and lighter assembly is possible.Those skilled in the art would recognize that, with the inventivemethod, layer-wise multi-material tubes and ducts can be enabled, whichpotentially provide novel mechanical & thermal properties. Moreover,surface texturing can be added to the interior surface of the metalducts and tubes for thermal transfer and pressure drop considerations.

The preferred carbon composite lay-up embodiment reduces productiontooling labor and material costs, for complex geometry parts. Theembodiments of the present invention obviate the need to bend, form andjoin tubes and ducts with costly complicated tools such as hydraulicmandrels. The unique conformable woven composite sleeve approach,described herein, with smart tooling is a simple solution to achallenging manufacturing process that often requires complex multipledegree of freedom robotic equipment and tooling. These complex roboticsystems have high initial equipment and programming development costs.

The embodiments thus use a totally new method than the currentmanufacturing methods for tube and duct. Low temperature techniques areused for layering material onto reusable SMP tooling, rather than hightemperature forming, joining processes. In some instances, hotdeposition metal forming processes or 3-D metal printing (Direct MetalLaser Melting) are used to produce similar parts. In other instances,wet lay-up of woven or spun carbon fiber on expendable 3-D mandrels canbe used. For the reasons indicated herein, however, it is believed thelow temperature deposition fabrication method provides more economicaland technical advantages than those other methods.

It is to be appreciated that the Detailed Description section, and notthe Summary and Abstract sections, is intended to be used to interpretthe claims. The Summary and Abstract sections may set forth one or morebut not all exemplary embodiments of the present invention ascontemplated by the inventor(s), and thus, are not intended to limit thepresent invention and the appended claims in any way.

What is claimed is:
 1. A method for fabricating aircraft engine externaltarget parts including complex geometries utilizing reusablereconfigurable shape memory polymer and conformable woven braided carbonfiber sleeves, the method comprising: providing a tubular 3-dimensionalreusable shape memory polymer mandrel assembly designed for a targetpart (Part C); providing conformable woven braided carbon fiber sleeves;providing a high temperature resin system formulated for carbon fiber;covering the shape memory polymer mandrel with a release agent; applyinga first layer of braided carbon fiber sleeve, to the shape memorypolymer mandrel, using the high temperature resin system for wetting andadhesion; using said resin system to overlay subsequent wetted layers tothe shape memory polymer mandrel, to create the final formed carboncomposite tube which replicates Part C; and, heating the shape memorypolymer mandrel until its temperature exceeds the glass transitiontemperature of the shape memory polymer, and the shape memory polymerbecomes compliant and elastic, and then, removing the elastic shapememory polymer and separating it from the newly formed Part C.
 2. Thefabrication method of claim 1, wherein, after separation, the shapememory polymer mandrel is cleaned, reheated, and placed in the toolingshape mold for re-forming and reuse.
 3. The fabrication method of claim1, wherein a subsequent higher temperature resin cure is performed, onPart C, after removal of the reusable polymer tool.
 4. The fabricationmethod of claim 1, wherein the high temperature resin system, formulatedfor carbon fiber, further comprises an optional vacuum bag system toremove excess resin.
 5. The fabrication method of claim 1, wherein PartC is a low temperature tube or duct.
 6. The fabrication method of claim1, wherein said fluid delivery parts are parts of an aerospace system.7. The fabrication method of claim 1, wherein said fluid delivery partsare external aircraft engine parts.
 8. The fabrication method of claim1, wherein said fluid delivery parts are parts of a turbine.
 9. Thefabrication method of claim 1, wherein the target parts include complexcylindrical and non-cylindrical tube geometries with non-uniformcross-sectional size (tapering).
 10. The fabrication method of claim 1,wherein the target part geometries include complex transitions fromcylindrical to rectangular cross-sectional shapes.
 11. The fabricationmethod of claim 1, wherein, depending on loads and boundary conditions,multiple local properly orientated braided sleeves are applied toincrease direction strength at high stress locations.
 12. Thefabrication method of claim 1, wherein Part C is a mounting bracket, fanhouse casing, or support.
 13. The fabrication method of claim 1, whereincomplex geometry tubes are formed before the thermoset resin is appliedand cured, thereby obviating post-bending and joining.
 14. Thefabrication method of claim 1, wherein the carbon fiber sleeves areproperly sized for the target part.